1. Field of the Invention
The present invention relates to solar panels for use in spacecraft and, more particularly, to solar cell assemblies which are the building blocks of solar panels and, still more particularly, to a unique robust construction of a solar cell assembly with significant cost and efficiency benefits.
2. Description of the Prior Art
Since the launch of Vanguard I on Mar. 17, 1958, photovoltaic panels have been the primary source of power generation for earth orbiting satellites in the United States, Europe and Japan. Space solar arrays and satellite power demands have increased since Vanguard's 1 watt array to the present power of nearly 20,000 watts at launch. These solar arrays are composed of thousands of individual solar cells configured in a parallel and series arrangement to satisfy the voltage and current requirements.
The solar cells ere of either silicon or gallium arsenide designs, usually rectangular in shape, of dimensions typically 4 cm by 6 cm, although sizes from 2 by 4 cm to a by 8 cm are in use. These cells are quite thin, 150 .mu.m to 200 .mu.m thick, fragile and easily damaged. To protect the cell during the process of building the solar array a subassembly of the cell, a thin glass cover and an interconnector is usually constructed. This provides significant strength protects the top surface of the cell and can be more easily handled. This is called a CIC or Cover Integrated Cell. This is the electrical building block for circuit construction.
An exploded view of a CIC is illustrated in FIG. 1, indicated by reference numeral 20. On the surface of a cell 22 are discrete contact pads 24 to which the electrical contact is made. The upper surface of the pads are silver to permit ease of soldering or welding. Interconnects 26 for electrically joining circuitry on the cell 22 to an adjoining cell 28 are either pure silver or silver plated kovar, molybdenum or Invar. These latter materials provide a better thermal expansion match to the cell material. Silver is soft and compliant, not requiring a perfect match. Since the satellites are subjected to eclipse cycles on orbit, the temperature excursion of the solar array for geostationary spacecraft is between -175.degree. C. and +60.degree. C. This temperature variation causes relative motion between cells due to the thermal expansion and contraction of the solar cells. The interconnect is designed to be flexible to accommodate this relative motion as indicated by the loops in the figure.
The cell 22 with interconnects 26 is then protected by bonding a cover glass 30 with a clear silicone adhesive 31. The adhesive has to be clear and able to survive the space environment. Dow Corning 93-500 is the accepted adhesive in Asia, Europe and USA. Although not illustrated, the adjoining cell 28 would be similarly protected.
The assembly of the CIC 20 is usually a custom operation because of the variety of cell sizes and requirements. This requires the hand operation of trained personnel with minimum reliance on automated equipment. The assembly flow follows the sequence described above: place the cell on a vacuum holding fixture, weld or solder the interconnect using an automatic welder or solder machine, apply the adhesive in measured quantity using an automation adhesive dispenser, position the glass against alignment pins, place in a small vacuum chamber to pull any trapped bubbles out of the cell the in an oven set at about 50.degree. C. to cure the adhesive. The CIC is then cleaned of any excess adhesive. Electrical and visual inspection is performed and the ClCs sorted by power output.
Spacecraft power demands continue to increase resulting in correspondingly increased requirements for the capability of the photovoltaic solar panels. As the demands continue to increase, the solar array becomes significantly larger and less able to use the lower efficiency and lower cost silicon solar cells and must consider the gallium arsenide family on high efficiency III-V (groups within the Periodic Table) solar cells. These solar cells are significantly more expensive than silicon based solar cells in the initial fabrication but also must be protected from reverse bias breakdown and severe power loss due to current limiting. The current limiting is typically due to shadowing of the solar cells or cell breakage that reduces the active area. The method to prevent the reverse bias is to incorporate a diode across the solar cell junction to provide the required current path around the cell.
Additionally, due to the high cost of the basic solar cell, every effort should be made to utilize the most active area of the blank cell. The assembly of these components together with a cover glass and stress relieved electrical interconnector in a robust, cost effective method is the subject of this disclosure.
A very common substrate for the manufacture of solar cells for spacecraft application for both silicon and gallium arsenide type solar cells is round semiconductor material of 100 mm diameter. Although other diameters, such as 75 mm, 125 mm and 150 mm are also considered or in use, the technique described herein is equally applicable. FIG. 2, for example, illustrates a 100 mm diameter solar cell wafer 32 with two typical rectangular solar cells 34, 36. The size of the maximum square configuration within the 100 mm circle is 70 mm by 70 mm, ignoring any edge clearance requirement. This square provides two cells 35 mm by 70 mm, each defining the solar cell total area of 2,450 mm.sup.2.
It was with knowledge of the foregoing state of the technology that the present invention has been conceived and is now reduced to practice.